This invention relates generally to the field of materials technologies, and more specifically to the field of component part fabrication and repair, and specifically to a process for fabricating or repairing a turbine component by applying layers of a material using a cold spray technique.
It is well known that the power and efficiency of operation of a gas turbine engine or a combined cycle power plant incorporating such a gas turbine engine may be increased by increasing the firing temperature in the combustion portions of the turbine. The demand for improved performance has resulted in advanced turbine designs wherein the peak combustion temperature may reach 1,400 degrees C. or more. Special materials are needed for components exposed to such temperatures. Nickel and cobalt based superalloy materials are now used for components in the hot gas flow path, such as combustor transition pieces and turbine rotating and stationary blades. Such superalloy materials are known in the art and include, for example, alloys 738, Mar M247 and CM 247LC. It is also known to coat a superalloy metal component with an insulating material to improve its ability to survive high operating temperatures in a combustion turbine environment. A ceramic top coat may be applied to a superalloy substrate structure with an intermediate metallic bond coat. Common ceramic insulating materials include yttria stabilized zirconia (YSZ), hafnia or scandia stabilized zirconia, and yttrium aluminum garnet (YAG). The bond coat layer provides oxidation resistance and improved adhesion for the thermal barrier coating layer. Common bond coat materials include MCrAlY and MCrAlRe, where M may be nickel, cobalt, iron or a mixture thereof. The metallic bond coat provides a level of thermal insulation, and in some applications may be used alone without an overlying ceramic layer.
Superalloy materials may be cast conventionally or as directionally solidified (DS) or single crystal (SC) material. Once the material is in the directionally solidified or single crystal structure, it is undesirable to subject the material to a combination of temperature and mechanical work that would result in its re-crystallization. A weld repair to a part formed of a directionally solidified or single crystal material may significantly degrade the material properties and operating performance of the part as a result of re-crystallization in the heat affected zone of the weld, as well as possible re-crystallization in other areas of the part occurring during any associated post-weld heat treatment.
FIG. 1 illustrates a top view of a prior art turbine blade 10 including a blade root 12, an airfoil portion 14 and a tip portion 16. The blade root 12 is designed to be inserted into and retained by a disc on a rotating shaft (not shown) of the turbine. The airfoil portion 14 is shaped to extract energy from combustion gases passing over the airfoil portion 14, thereby imparting rotating mechanical energy to the turbine shaft. A thermal barrier coating, such as described above, may be applied to a portion of the airfoil portion 14. Airfoil portion 14 may be designed to include one or more cooling passages formed below the surface of the airfoil for the passage of cooling air necessary to ensure the integrity of the blade material in the hot combustion gas environment. During the manufacturing process, such cooling passages are drilled or cast to extend from the edge of the blade tip portion 16. These openings must then be sealed during the fabrication process in order to assure the proper flow of the cooling air within the blade 10. If the size of the opening is sufficiently small, it may be sealed by a weld plug 18. For larger openings, it may be necessary to cover the opening with a cap, such one or more plates 20, in order to seal the opening. U.S. Pat. No. 4,073,599 issued on Feb. 14, 1978, to Allen et al. describes such a blade tip closure design. Plates 20 are mechanically restrained by the structure of the blade tip 16 and are held in position and sealed by one or more brazed joints 22.
It is known that turbine blades 10 may develop one or more cracks 24 near the tip 16 of the blade 10 due to low cycle fatigue stresses imparted on the tip 16 during the operation of the turbine. The turbine blade 10 must be removed from service and/or repaired before a crack 24 extends beyond a critical dimension in order to prevent catastrophic failure of the blade and turbine. It can be appreciated that a crack 24 may be repaired by removing the material adjacent to the crack 24 to form a crack repair volume, and then filling the crack repair volume with weld metal. However, the presence of braze joint 22 can complicate the repair process, since weld integrity is adversely affected when applied over a braze material.
The repair of turbine component parts is complicated by the difficulties associated with welding, in general, and the welding of superalloy materials, in particular. Welding operations are further complicated by the presence of contaminants on the surface of the part to be repaired, and the proximity of braze material in the vicinity of the weld. Post-weld heat treatment adds time and cost to the repair, and it may further degrade the properties of directionally solidified or single crystal base materials. Thus, an improved process is needed for repairing turbine component parts.
The present invention utilizes a,cold spray material deposition process in lieu of welding to fabricate or repair a component part. By depositing repair material with a cold spray process, the re-crystallization of directionally solidified and single crystal base materials is prevented, and the need for a high temperature heat treatment is eliminated. Repairs employing the cold spray deposition of repair material may be made directly over brazed material. Insert materials may be joined to a part by forming a joint from cold sprayed material. The immediate area of a repair may be cleaned of contaminants during a cold spray material deposition step by the grit blasting effect of a cold spray pattern halo.
Accordingly, a process for repairing a component part is described herein as including the steps of: identifying a discontinuity in a part surface; excavating material from the part surface proximate the discontinuity to form a repair surface; and directing particles of a repair material toward the part at a velocity sufficiently high to cause the particles to deform and to adhere to the repair surface.
A process for repairing a turbine is described herein as including the steps of: at least partially disassembling a turbine to provide access to a part having an area to be repaired without removing the part from the turbine; directing particles of a repair material toward the area to be repaired at a velocity sufficiently high to cause the particles to deform and to adhere to the surface of the area to be repaired; and re-assembling the turbine.
A process for repairing a turbine blade having a crack in a tip portion is described, the tip portion including a plate joined to the tip portion by a braze joint, the crack being proximate the braze joint, the process comprising the steps of: preparing the surface of the tip portion proximate the crack to form a repair surface, the repair surface extending to include at least a portion of the braze joint; and directing particles of a repair material toward the repair surface at a velocity sufficiently high to cause the particles to deform and to adhere to the repair surface.
A metals joining process is described as including the steps of: forming a first joining surface on a first part comprising a first metal; forming a second joining surface on a second part comprising a second metal; positioning the first joining surface proximate the second joining surface; and directing particles of a joining material toward the first and the second joining surfaces at a velocity sufficiently high to cause the particles to deform and to adhere to the first and the second joining surfaces to form a joint there between.
A process for repairing a part is described herein as including the steps of: identifying a discontinuity in a surface of a part; directing a pattern of particles of a repair material toward the surface of the part, the pattern having a center area and a halo area surrounding the center area; wherein the speed of the particles in a direction perpendicular to the surface of the part is sufficiently high in the center area to cause the particles to deform and to adhere to the surface of the part; and wherein the speed of the particles in a direction perpendicular to the surface of the part is not sufficiently high in the halo area to cause the particles to deform and to adhere to the surface of the part but is sufficiently high to remove contaminants from the surface of the part; moving the pattern of particles across the discontinuity so that particles in the halo area first clean the surface of the part proximate the discontinuity and particles in the center area then adhere to the surface of the part to repair the discontinuity.
A process for forming a desired geometry on a part, the process comprising the steps of: forming a part having a geometry other than a desired geometry; directing particles of a material toward the part at a velocity sufficiently high to cause the particles to deform and to adhere to the surface of the part; and continuing the step of directing particles until the desired geometry is formed.
In a further embodiment, a process for repairing a component part is described, the process including the steps of: identifying a discontinuity producing an unacceptable stress rise in a part; directing particles of a repair material toward the discontinuity at a velocity sufficiently high to cause the particles to deform and to adhere to the surface of the discontinuity; and continuing the step of directing particles until the discontinuity has been filled with repair material to an extent necessary to reduce the stress rise to a desired level.